Power plants – Reaction motor – Solid and fluid propellant
Reexamination Certificate
1999-07-21
2004-02-03
Miller, Edward A. (Department: 3641)
Power plants
Reaction motor
Solid and fluid propellant
C149S018000, C149S109400
Reexamination Certificate
active
06684624
ABSTRACT:
FIELD OF THE INVENTION
The present invention relates generally to the field of propellants suitable for use in hybrid rockets, and more particularly to propellants and a method of selecting propellants that exhibit high regression rates.
BACKGROUND OF THE INVENTION
Two basic types of chemical rocket propulsion systems are widely used in the rocket industry; namely, liquid systems and solid propellant systems. In a liquid system, liquid oxidizer and liquid fuel are fed at high pressure to a combustion chamber where they mix and react producing high temperature, high pressure gases which exhaust through a converging-diverging nozzle producing thrust. The mixing of reactants requires a high performance pressurization system for the fuel and oxidizer which must often operate in a cryogenic environment at extreme pressures and mass flow rates. The injection system and combustion chamber require exotic materials, complex systems for cooling, and very high precision manufacturing techniques. All of these factors contribute to a high cost.
Solid propellant systems do not require the complex and expensive machinery of liquid systems. Nevertheless, solid systems are complicated, and are subject to the difficulties of producing crack-free, repeatable, fuel grains, and by the need to transport and handle explosive materials. In a manufacturing process that requires extreme safety precautions, solid fuel and oxidizer are intimately mixed and allowed to cure inside the rocket case producing an explosive fuel with roughly the consistency of plastic or hard rubber. Fuel grains which contain cracks present a risk of explosive failure of the vehicle and must be rejected, driving up the cost of manufacture. Upon ignition the solid fuel burns uninterrupted until all the fuel is exhausted.
An alternative chemical rocket which has been known since the 1930's is the hybrid propulsion system. In the hybrid design one propellant is stored in the solid phase while the other is stored in the liquid phase. Thus the hybrid lies somewhere between the two basic chemical rocket designs just described. In most hybrid propulsion applications, the solid is the fuel and the liquid is the oxidizer. Reverse hybrids with the fuel in the liquid phase and oxidizer in the solid phase are also feasible and the present invention described below can be applied equally well to both types of hybrid systems.
A large variety of fuels, including trash and wood, have been considered for hybrid rockets but the most conventional fuel materials are polymers such as Plexiglas (polymethyl methacrylate) (PMMA), high density polyethylene (HDPE), hydroxyl terminated polybutadiene (HTPB), and the like. Typical oxidizers that are frequently used in hybrid rockets are liquid oxygen, hydrogen peroxide, nitrogen tetroxide, nitrous oxide and occasionally fluorine. With respect to the last point, the fuel combinations used for hybrids are similar in their chemical properties and energy densities to the fuels used in hydrocarbon fueled liquid rocket systems. Thus, in terms of exhaust velocity and specific impulse, the hybrid system is a closer relative to a liquid system than to a solid system. Solid rockets tend to use lower energy oxidizers and consequently they produce lower specific impulse.
In addition to having a higher specific impulse, some of the advantages of the hybrid rocket over the solid fuel rocket are:
The hybrid allows for thrust termination, restart and throttling capabilities,
The hybrid design lends itself to safe manufacturing, transportation and operation.
Hybrid motors are inherently immune to explosion,
The safety and simplicity of the hybrid leads to lower development costs for new systems and very likely lower operational costs,
The combustion products are generally very benign producing lower environmental impact.
The main advantages of the hybrid over the liquid rocket include:
Lower development and operating costs (life cycle costs),
Lower fire and explosion hazards,
Less complex design and potentially higher reliability.
The hybrid allows the addition of energetic solid components, such as aluminum or beryllium to the fuel.
A schematic of a typical hybrid propulsion system
10
with a pressurized oxidizer feed system is shown in FIG.
1
. The feed system is comprised of a pressurization tank
12
that holds an inert gas at high pressure (such as Helium, Argon or Nitrogen), a valve (not shown) to pressurize the oxidizer tank
14
, a main valve
16
to turn on the flow of oxidizer and an injection system 18. Alternatively, the gas pressurization system can be replaced with a turbopump. The other major components are the combustion chamber
20
which contains the solid fuel
22
and the nozzle assembly
24
.
A sketch of the flame configuration in a single port hybrid rocket combustion chamber
30
is shown in FIG.
2
. The single port combustion chamber
30
generally includes a pre-combustion chamber region
31
at the front end, a post-combustion chamber region
32
at the opposite end, and an elongated single port
33
extending between the ends. The oxidizer in the liquid phase is injected into the combustion chamber at pre-combustion chamber region
31
. The injected oxidizer is gasified and flows axially along the port
33
, forming a boundary layer
34
over the solid fuel
22
. The boundary layer
34
is usually turbulent in nature over a large portion of the length of the port. Within the boundary layer
34
there exists a turbulent diffusion flame
36
which extends over the entire length of the fuel. The thickness of the flame is generally very small compared to the boundary layer thickness. The heat generated in the flame, which is located approximately 20-30% of the boundary layer thickness above the fuel surface, is transferred to the wall mainly by convection. Some heat is also transferred by radiation but this is usually relatively small compared to the convective heat transfer. In the conventional hybrid system depicted in
FIG. 2
, the wall heat flux evaporates the (generally polymeric) solid fuel and the fuel vapor is transported to the flame where it reacts with the oxidizer which is transported from the free stream by turbulent diffusion mechanisms. The unburned fuel that travels beneath the flame, the unburned oxidizer in the free stream, and the flame combustion products mix and further react in the post combustion chamber
32
. The degree to which fuel and oxidizer are able to fully mix and react before exhausting through the nozzle
24
determines the combustion efficiency of the motor. The hot gases expand through a convergent-divergent nozzle
24
to deliver the required thrust.
It is important to note that, even though the geometry of a hybrid motor is similar to a solid motor, the combustion scheme is vastly different. In a solid rocket, the oxidizer and fuel are both stored in the solid phase next to each other for heterogeneous fuels and within the same fuel molecule for double base fuels. Consequently, the solid combustion takes place in a deflagration (premixed) flame that is closer to the surface compared to the hybrid diffusion flame. Also, in solid fuel systems there exists some heterogeneous phase (solid-solid, solid-gas) reactions at the surface. In short, the burning rate of a solid rocket is determined by the rate of homogeneous (gas phase) and heterogeneous chemical reactions.
In a hybrid system or motor, the burning rate is limited by the heat transfer from the relatively remote flame to the burning surface of the fuel. One of the physical phenomena that limits the burning rate in a hybrid motor is the so-called blocking effect that is caused by the high velocity injection of the vaporizing fuel into the gas stream. This difference in the combustion scheme of a hybrid motor significantly alters the burning rate characteristics compared to a solid rocket. Blocking can be explained as follows. Increasing the heat transfer to the fuel causes the evaporative mass transfer from the liquid-gas interface to increase. But the increased blowing from the surface reduces the temp
Altman David
Cantwell Brian J.
Karabeyoglu M. Arif
Dorsey & Whitney LLP
Miller Edward A.
The Board of Trustees of the Leland Stanford Junior University
LandOfFree
High regression rate hybrid rocket propellants does not yet have a rating. At this time, there are no reviews or comments for this patent.
If you have personal experience with High regression rate hybrid rocket propellants, we encourage you to share that experience with our LandOfFree.com community. Your opinion is very important and High regression rate hybrid rocket propellants will most certainly appreciate the feedback.
Profile ID: LFUS-PAI-O-3327952